US6438958B1 - Apparatus for reducing heat load in combustor panels - Google Patents

Apparatus for reducing heat load in combustor panels Download PDF

Info

Publication number
US6438958B1
US6438958B1 US09/513,943 US51394300A US6438958B1 US 6438958 B1 US6438958 B1 US 6438958B1 US 51394300 A US51394300 A US 51394300A US 6438958 B1 US6438958 B1 US 6438958B1
Authority
US
United States
Prior art keywords
liner
panel
panels
combustor
thermal barrier
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
US09/513,943
Inventor
Timothy P. McCaffrey
Frank A. Lastrina
Joseph D. Monty
David E. Hrencecin
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
General Electric Co
Original Assignee
General Electric Co
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by General Electric Co filed Critical General Electric Co
Priority to US09/513,943 priority Critical patent/US6438958B1/en
Assigned to GENERAL ELECTRIC COMPANY reassignment GENERAL ELECTRIC COMPANY ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: HRENCECIN, DAVID E., LASTRINA, FRANK A., MC CAFFREY, TIMOTHY P., MONTY, JOSEPH D.
Priority to CA002337311A priority patent/CA2337311C/en
Priority to BRPI0100759-9A priority patent/BR0100759B1/en
Priority to JP2001051646A priority patent/JP2001248839A/en
Priority to EP01301843A priority patent/EP1132686B1/en
Priority to DE60122954T priority patent/DE60122954T2/en
Priority to US10/192,576 priority patent/US6519850B2/en
Publication of US6438958B1 publication Critical patent/US6438958B1/en
Application granted granted Critical
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/002—Wall structures
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/06—Arrangement of apertures along the flame tube
    • F23R3/08—Arrangement of apertures along the flame tube between annular flame tube sections, e.g. flame tubes with telescopic sections
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/00005—Preventing fatigue failures or reducing mechanical stress in gas turbine components
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/49—Method of mechanical manufacture
    • Y10T29/49346—Rocket or jet device making

Definitions

  • This invention relates generally to turbine engines, and, more particularly, to slot cooled ring combustors for turbine engines.
  • a turbine engine includes a compressor for compressing air which is mixed with a fuel and channeled to a combustor wherein the mixture is ignited for generating hot combustion gases.
  • the combustion gases are channeled to a turbine, which extracts energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator.
  • Increased efficiency in gas turbine engines is accomplished at least in part by an increase in the operating temperature of the combustor.
  • a limitation on the operating combustor temperature is a temperature limitation of combustor liner material.
  • Thin film convection cooling can be used to cool a combustor liner. With such cooling, a protective film boundary of cool air flows along an inner surface of the liner. The cool air flowing along the combustor liner inner surface forms a is protective boundary between the liner and the hot gases, and insulates the liner from hot combustion gases. See, for example, U.S. Pat. No. 4,259,842. Even with such cooling, however, the liner materials absorb heat. Over time, thermal creep and low cycle fatigue increase in the liner.
  • a thermal barrier coating also can be applied to inner surfaces of the combustor liner for providing thermal insulation against combustion gases.
  • Thermal barrier coatings reduce an amount of cooling air required for a given combustion gas temperature, or allow an increase in a combustion gas temperature for increasing efficiency of the engine. See, for example, U.S. Pat. No. 5,960,632.
  • the thermal barrier coating is applied uniformly across the combustor liner with a thickness of 0.01 inches or less. Such a uniform thickness prevents the thermal barrier coating from undesirably building-up to potentially obstruct the flow of cooling air.
  • the combustor liner materials still absorb heat, and thus, combustor assemblies are still subjected to thermal strains including creep and low cycle fatigue.
  • a combustor in an exemplary embodiment, includes a combustor liner s with a thermal barrier material that has a thickness selected to minimize heat absorption.
  • the combustor includes a combustion zone formed by annular outer and inner supporting members and respective inner and outer liners.
  • the inner and outer liners each include a series of panels and a plurality of cooling slots. The panels are arranged in steps relative to one another and form a stepped combustor liner surface.
  • the plurality of cooling slots are formed by overhanging portions of the inner and outer liner panels.
  • At least one portion of the combustor liner has a thermal barrier material with a thickness greater than 0.01 inches.
  • at least the outer and inner liner panels adjacent an inlet of the combustor have a thermal barrier material with a thickness greater than 0.01 inches.
  • the combustor liner material absorbs less heat, and therefore, at present day operating temperatures, the combustor may be operated at higher temperatures. Because the operating temperature is reduced, low cycle fatigue within the combustor is also reduced which, in turn, extends an operating life cycle of the combustor assembly.
  • FIG. 1 is a schematic illustration of a gas turbine engine
  • FIG. 2 is partial cross-sectional view of a combustor assembly used with the gas turbine engine shown in FIG. 1;
  • FIG. 3 is an enlarged view of a portion of the combustor assembly shown in FIG. 2 taken along area 2 .
  • FIG. 1 is a schematic illustration of a gas turbine engine 10 including a low pressure compressor 12 , a high pressure compressor 14 , and a combustor assembly 16 .
  • Engine 10 also includes a high pressure turbine 18 , and a low pressure turbine 20 .
  • Compressor 12 and turbine 20 are coupled by a first shaft 24
  • compressor 14 and turbine 18 are coupled by a second shaft 26 .
  • engine 10 is a CF34-3A/-3B engine available from General Electric Aircraft Engines, Cincinnati, Ohio.
  • Compressed air is then delivered to combustor assembly 16 where it is mixed with fuel an ignited.
  • the combustion gases are channeled from combustor 16 to drive turbines 18 and 20 .
  • FIG. 2 is a partial cross-sectional view of combustor assembly 16 including a combustor 30 and a fuel injector 32 .
  • FIG. 3 is an enlarged view of a portion of combustor assembly 16 .
  • Fuel injector 32 is attached to an inlet 34 of combustor assembly 16 and injects atomized fuel into a combustion zone 36 of combustor 30 to form an air-fuel mixture.
  • An igniter or cross-fire tube ignites the air-fuel mixture downstream of fuel injector 32 , and combustion gases exit combustor assembly 16 through an outlet turbine nozzle 38 that directs high energy combustion gases towards a row of turbine blades or buckets (not shown).
  • the combustion gases rotate a turbine wheel (not shown) that delivers rotational energy to a compressor (not shown), powers a load, and/or is converted into thrust.
  • Combustion zone 36 is formed by annular, radially outer and radially inner supporting members (not shown) and a combustor liner 40 .
  • Combustor liner 40 shields the outer and inner supporting members from the heat generated within combustion zone 36 and includes an outer liner 50 and an inner liner 52 .
  • Outer liner 50 and inner liner 52 are annular and connect together to define combustion zone 36 .
  • Combustion zone 36 extends from combustor inlet 34 to outlet turbine nozzle 38 .
  • Outer and inner liners 50 and 52 each include a plurality of panels 54 which include a series of steps 56 , each of which form a distinct portion of combustor liner 40 .
  • Outer liner 50 and inner liner 52 each include a cowl 60 and 62 , respectively, and a first panel 64 and 66 , respectively.
  • Inner cowl 60 and outer cowl 62 are positioned adjacent combustor inlet 34 and extend from combustor inlet 34 to first panels 64 and 66 , respectively.
  • First panels 64 and 66 are connected serially downstream from cowls 60 and 62 , respectively, and each are connected between cowls 60 and 62 , respectively, and additional outer and inner liner panels 54 .
  • outer liner 50 and inner liner 52 each include nine panels 54 .
  • Each combustor panel 54 includes a combustor liner surface 70 , an exterior surface 72 , and an overhang portion 74 .
  • Combustor liner surface 70 extends from combustor inlet 34 to outlet turbine nozzle 38 .
  • Combustor liner surface 70 and exterior surface 72 are connected together at overhang portion 74 and form a rear facing edge 76 .
  • a plurality of air cooling slots 78 separate adjacent combustor panels 54 .
  • Air cooling slots 78 include openings 80 to receive air from an air plenum (not shown) and form a thin protective boundary of air between high temperature combustion gases and combustor liner surface 70 , as well as providing for convective cooling of combustor liner 40 . Air flows from openings 80 through slots 78 formed between combustor liner surface 70 and a bottom surface 82 of combustor liner overhang portions 74 .
  • a layer 90 of thermal barrier material is applied on combustor liner surface 70 and extends from overhang portion 74 to overhang portion 74 of each step 54 .
  • Thermal barrier material further insulates combustor liner surface 70 from high temperature combustion gases.
  • thermal barrier material is commercially available from Englehart Industries, Wilmington Mass.
  • Thermal barrier material is applied to combustor liner surface 70 over each combustor panel 54 disposed between combustor inlet 34 and combustor outlet turbine nozzle 38 .
  • Thermal barrier material is applied such that layer 90 has a thickness T 1 greater than 0.01 inches extending over at least a portion 96 of combustor liner surface 70 .
  • portion 96 includes only outer liner first panel 64 and inner liner first panel 66 and any remaining panels 54 have a layer 90 of thermal barrier material with a thickness T 2 of 0.01 inches or less.
  • portion 96 includes outer and inner liner first panels 64 and 66 , and at least one other outer liner panel 54 and inner liner panel 54 , and any remaining panels 54 have thermal barrier material with thickness T 2 is 0.01 inches or less.
  • combustor liner surface 70 thermal barrier material is applied such that layer 90 extends over all combustor panels 54 between combustor inlet 34 and combustor outlet turbine nozzle 38 and has thickness T 1 greater than 0.01 inches.
  • Thickness T 1 is measured from combustor liner surface 70 to a top surface 98 of layer 90 .
  • layer 90 extends over portion 96 and has thickness T 1 approximately twice thickness T 2 of thermal barrier material extending over panels 54 not in portion 96 .
  • thermal barrier material thickness T 1 is between 0.20 and 0.35 inches and thickness T 2 is 0.01 inches or less. In a further embodiment, thermal barrier material thickness T 1 is approximately 0.20 inches.
  • portion 96 has thermal barrier material with a thickness T 1 less heat is absorbed into panels 54 within combustor portion 96 and an operating temperature of combustor 30 is lowered, thus reducing an amount of thermal strains within combustor assembly 16 .
  • the above-described combustor assembly is cost-effective and highly reliable.
  • the combustor assembly includes a thermal barrier material having a thickness greater than 0.01 inches covering at least a portion of the combustor liner. As a result, the combustor liner absorbs less heat, and therefore, the combustor may be operated at higher temperatures. Because the operating temperature is reduced, low cycle fatigue within the combustor is reduced, which in turn, extends an operating life cycle for the combustor assembly.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Gas Burners (AREA)
  • Solid Fuels And Fuel-Associated Substances (AREA)

Abstract

A combustor liner has a stepped combustor liner surface defining a combustion zone and an overhang portion forming an air cooling slot. A layer of thermal barrier material is applied to the combustor liner such that at least one portion of the combustor liner receives a layer of thermal barrier material with a thickness greater than 0.01 inches. Thus, the combustor liner absorbs less heat, and the combustor may operate at higher temperatures. As a result, low cycle fatigue and thermal creep are reduced within the combustor and the life cycle for the combustor is extended.

Description

BACKGROUND OF THE INVENTION
This invention relates generally to turbine engines, and, more particularly, to slot cooled ring combustors for turbine engines.
A turbine engine includes a compressor for compressing air which is mixed with a fuel and channeled to a combustor wherein the mixture is ignited for generating hot combustion gases. The combustion gases are channeled to a turbine, which extracts energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator. Increased efficiency in gas turbine engines is accomplished at least in part by an increase in the operating temperature of the combustor. A limitation on the operating combustor temperature is a temperature limitation of combustor liner material.
Thin film convection cooling can be used to cool a combustor liner. With such cooling, a protective film boundary of cool air flows along an inner surface of the liner. The cool air flowing along the combustor liner inner surface forms a is protective boundary between the liner and the hot gases, and insulates the liner from hot combustion gases. See, for example, U.S. Pat. No. 4,259,842. Even with such cooling, however, the liner materials absorb heat. Over time, thermal creep and low cycle fatigue increase in the liner.
A thermal barrier coating also can be applied to inner surfaces of the combustor liner for providing thermal insulation against combustion gases. Thermal barrier coatings reduce an amount of cooling air required for a given combustion gas temperature, or allow an increase in a combustion gas temperature for increasing efficiency of the engine. See, for example, U.S. Pat. No. 5,960,632. Typically the thermal barrier coating is applied uniformly across the combustor liner with a thickness of 0.01 inches or less. Such a uniform thickness prevents the thermal barrier coating from undesirably building-up to potentially obstruct the flow of cooling air. However, the combustor liner materials still absorb heat, and thus, combustor assemblies are still subjected to thermal strains including creep and low cycle fatigue.
BRIEF SUMMARY OF THE INVENTION
In an exemplary embodiment, a combustor includes a combustor liner s with a thermal barrier material that has a thickness selected to minimize heat absorption. In the exemplary embodiment, the combustor includes a combustion zone formed by annular outer and inner supporting members and respective inner and outer liners. The inner and outer liners each include a series of panels and a plurality of cooling slots. The panels are arranged in steps relative to one another and form a stepped combustor liner surface. The plurality of cooling slots are formed by overhanging portions of the inner and outer liner panels. At least one portion of the combustor liner has a thermal barrier material with a thickness greater than 0.01 inches. In the exemplary embodiment, at least the outer and inner liner panels adjacent an inlet of the combustor have a thermal barrier material with a thickness greater than 0.01 inches.
As a result of the additional thickness of thermal barrier material applied to at least a portion of the combustor liner, the combustor liner material absorbs less heat, and therefore, at present day operating temperatures, the combustor may be operated at higher temperatures. Because the operating temperature is reduced, low cycle fatigue within the combustor is also reduced which, in turn, extends an operating life cycle of the combustor assembly.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a schematic illustration of a gas turbine engine;
FIG. 2 is partial cross-sectional view of a combustor assembly used with the gas turbine engine shown in FIG. 1; and
FIG. 3 is an enlarged view of a portion of the combustor assembly shown in FIG. 2 taken along area 2.
DETAILED DESCRIPTION OF THE INVENTION
FIG. 1 is a schematic illustration of a gas turbine engine 10 including a low pressure compressor 12, a high pressure compressor 14, and a combustor assembly 16. Engine 10 also includes a high pressure turbine 18, and a low pressure turbine 20. Compressor 12 and turbine 20 are coupled by a first shaft 24, and compressor 14 and turbine 18 are coupled by a second shaft 26. In one embodiment, engine 10 is a CF34-3A/-3B engine available from General Electric Aircraft Engines, Cincinnati, Ohio.
In operation, air flows through low pressure compressor 12 from an inlet side 28 of engine 10 and compressed air is supplied from low pressure compressor 12 to high pressure compressor 14. Compressed air is then delivered to combustor assembly 16 where it is mixed with fuel an ignited. The combustion gases are channeled from combustor 16 to drive turbines 18 and 20.
FIG. 2 is a partial cross-sectional view of combustor assembly 16 including a combustor 30 and a fuel injector 32. FIG. 3 is an enlarged view of a portion of combustor assembly 16. Fuel injector 32 is attached to an inlet 34 of combustor assembly 16 and injects atomized fuel into a combustion zone 36 of combustor 30 to form an air-fuel mixture. An igniter or cross-fire tube (not shown) ignites the air-fuel mixture downstream of fuel injector 32, and combustion gases exit combustor assembly 16 through an outlet turbine nozzle 38 that directs high energy combustion gases towards a row of turbine blades or buckets (not shown). The combustion gases rotate a turbine wheel (not shown) that delivers rotational energy to a compressor (not shown), powers a load, and/or is converted into thrust.
Combustion zone 36 is formed by annular, radially outer and radially inner supporting members (not shown) and a combustor liner 40. Combustor liner 40 shields the outer and inner supporting members from the heat generated within combustion zone 36 and includes an outer liner 50 and an inner liner 52. Outer liner 50 and inner liner 52 are annular and connect together to define combustion zone 36. Combustion zone 36 extends from combustor inlet 34 to outlet turbine nozzle 38. Outer and inner liners 50 and 52 each include a plurality of panels 54 which include a series of steps 56, each of which form a distinct portion of combustor liner 40.
Panels 54 are connected serially. Outer liner 50 and inner liner 52 each include a cowl 60 and 62, respectively, and a first panel 64 and 66, respectively. Inner cowl 60 and outer cowl 62 are positioned adjacent combustor inlet 34 and extend from combustor inlet 34 to first panels 64 and 66, respectively. First panels 64 and 66 are connected serially downstream from cowls 60 and 62, respectively, and each are connected between cowls 60 and 62, respectively, and additional outer and inner liner panels 54. In one embodiment, outer liner 50 and inner liner 52 each include nine panels 54.
Each combustor panel 54 includes a combustor liner surface 70, an exterior surface 72, and an overhang portion 74. Combustor liner surface 70, extends from combustor inlet 34 to outlet turbine nozzle 38. Combustor liner surface 70 and exterior surface 72 are connected together at overhang portion 74 and form a rear facing edge 76. A plurality of air cooling slots 78 separate adjacent combustor panels 54.
Air cooling slots 78 include openings 80 to receive air from an air plenum (not shown) and form a thin protective boundary of air between high temperature combustion gases and combustor liner surface 70, as well as providing for convective cooling of combustor liner 40. Air flows from openings 80 through slots 78 formed between combustor liner surface 70 and a bottom surface 82 of combustor liner overhang portions 74.
A layer 90 of thermal barrier material is applied on combustor liner surface 70 and extends from overhang portion 74 to overhang portion 74 of each step 54. Thermal barrier material further insulates combustor liner surface 70 from high temperature combustion gases. In an exemplary embodiment, thermal barrier material is commercially available from Englehart Industries, Wilmington Mass. Thermal barrier material is applied to combustor liner surface 70 over each combustor panel 54 disposed between combustor inlet 34 and combustor outlet turbine nozzle 38. Thermal barrier material is applied such that layer 90 has a thickness T1 greater than 0.01 inches extending over at least a portion 96 of combustor liner surface 70. In one embodiment, portion 96 includes only outer liner first panel 64 and inner liner first panel 66 and any remaining panels 54 have a layer 90 of thermal barrier material with a thickness T2 of 0.01 inches or less. In another embodiment, portion 96 includes outer and inner liner first panels 64 and 66, and at least one other outer liner panel 54 and inner liner panel 54, and any remaining panels 54 have thermal barrier material with thickness T2 is 0.01 inches or less. In a further embodiment, combustor liner surface 70 thermal barrier material is applied such that layer 90 extends over all combustor panels 54 between combustor inlet 34 and combustor outlet turbine nozzle 38 and has thickness T1 greater than 0.01 inches.
Thickness T1 is measured from combustor liner surface 70 to a top surface 98 of layer 90. In one embodiment, layer 90 extends over portion 96 and has thickness T1 approximately twice thickness T2 of thermal barrier material extending over panels 54 not in portion 96. In another embodiment, thermal barrier material thickness T1 is between 0.20 and 0.35 inches and thickness T2 is 0.01 inches or less. In a further embodiment, thermal barrier material thickness T1 is approximately 0.20 inches.
During operation, as atomized fuel is injecting into combustion zone 36 and ignited, heat is generated within zone 36. Air enters combustion zone 36 through cooling slots 78 and forms a thin protective boundary of air along combustor liner surface 70. Combustor inner and outer liners 52 and 50 shield gas turbine engine 10 from heat generated within combustion zone 36. Because portion 96 has thermal barrier material with a thickness T1 less heat is absorbed into panels 54 within combustor portion 96 and an operating temperature of combustor 30 is lowered, thus reducing an amount of thermal strains within combustor assembly 16.
The above-described combustor assembly is cost-effective and highly reliable. The combustor assembly includes a thermal barrier material having a thickness greater than 0.01 inches covering at least a portion of the combustor liner. As a result, the combustor liner absorbs less heat, and therefore, the combustor may be operated at higher temperatures. Because the operating temperature is reduced, low cycle fatigue within the combustor is reduced, which in turn, extends an operating life cycle for the combustor assembly.
While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.

Claims (12)

What is claimed is:
1. A liner comprising a series of panels arranged in steps relative to one another, said panels separated by a series of cooling slots formed by overhanging portions of said series of panels, each said panel comprising a liner surface and an exterior surface, said panels having a layer of thermal barrier material applied only against said panel liner surface such that at least one of said panels has a layer of thermal barrier material having a thickness of greater than 0.01 inches, and such that an upstream edge of the thermal barrier material applied to each panel is downstream an adjacent upstream panel overhanging portion.
2. A liner in accordance with claim 1 further comprising at least one panel having a layer of thermal barrier material having a thickness at least twice that of a layer of thermal barrier material of said remaining panels.
3. A liner in accordance with claim 1 wherein said series of panels comprises a first panel, and a second panel, said first panel connected serially to said first panel.
4. A liner in accordance with claim 3 wherein said first panel has a layer of thermal barrier material having a thickness greater than 0.01 inches.
5. A liner in accordance with claim 4 wherein said first panel has a layer of thermal barrier material having a thickness at least twice that of a layer of thermal barrier material covering said second panel.
6. A combustor comprising:
an inlet;
an outlet;
a radially inner liner extending between said inlet and said outlet, said inner liner comprising a series of panels arranged in steps relative to one another, each said panel comprising a liner and a exterior surface, said inner liner panels separated by a series of cooling slots formed by overhanging portions of said inner liner panels, said inner liner panels having a layer of thermal barrier material, and a radially outer liner extending between said inlet and said outlet, said inner liner and said outer liner connected to form a combustion zone, said outer liner comprising a series of panels arranged in steps relative to one another, each said panel comprising a liner and an exterior surface, said outer liner panels separated by a series of cooling slots formed by overhanging portions of said outer liner panels, said outer liner panels having a layer of thermal barrier material, at least one of said inner liner panels and said outer liner panels having a layer of thermal barrier material applied only against said panel liner surface and having a thickness greater than 0.01 inches, greater than 0.01 inches, such that an upstream edge of the thermal barrier material applied to each panel liner surface is downstream an adjacent upstream panel exterior surface.
7. A combustor in accordance with claim 6 wherein said inner liner series of panels comprise a first panel and a second panel, said inner liner second panel serially connected downstream from said inner liner first panel, said outer liner series of panels comprise a first panel and a second panel, said outer liner first panel adjacent said combustor inlet, said outer liner second panel connected serially downstream from said outer liner first panel.
8. A combustor in accordance with claim 7 wherein at least one of said inner liner first panel and said outer liner first panel has a layer of thermal barrier material having a thickness greater than 0.01 inches.
9. A combustor in accordance with claim 7 wherein at least one of said inner liner panels and said outer liner panels has a layer of thermal barrier material having a thickness of approximately 0.02 inches.
10. A combustor in accordance with claim 9 wherein at least one of said inner liner first panel and said outer liner first panel has a layer of thermal barrier material having a thickness of approximately 0.02 inches.
11. A combustor in accordance with claim 8 wherein said inner and said outer liner first panels have a layer of thermal barrier material having a thickness greater than 0.01 inches.
12. A combustor in accordance with claim 11 wherein said inner and said outer liner first panels have a layer of thermal barrier material having a thickness of approximately 0.02 inches.
US09/513,943 2000-02-28 2000-02-28 Apparatus for reducing heat load in combustor panels Expired - Lifetime US6438958B1 (en)

Priority Applications (7)

Application Number Priority Date Filing Date Title
US09/513,943 US6438958B1 (en) 2000-02-28 2000-02-28 Apparatus for reducing heat load in combustor panels
CA002337311A CA2337311C (en) 2000-02-28 2001-02-15 Methods and apparatus for reducing heat load in combustor panels
BRPI0100759-9A BR0100759B1 (en) 2000-02-28 2001-02-23 method for manufacturing a combustor for a turbine engine, as well as coating and combustor.
JP2001051646A JP2001248839A (en) 2000-02-28 2001-02-27 Method and apparatus for reducing heat load in combustor panel
EP01301843A EP1132686B1 (en) 2000-02-28 2001-02-28 Methods and apparatus for reducing heat load in combustor panels
DE60122954T DE60122954T2 (en) 2000-02-28 2001-02-28 Method and device for reducing the heat load on a combustion chamber wall
US10/192,576 US6519850B2 (en) 2000-02-28 2002-07-10 Methods for reducing heat load in combustor panels

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US09/513,943 US6438958B1 (en) 2000-02-28 2000-02-28 Apparatus for reducing heat load in combustor panels

Related Child Applications (1)

Application Number Title Priority Date Filing Date
US10/192,576 Division US6519850B2 (en) 2000-02-28 2002-07-10 Methods for reducing heat load in combustor panels

Publications (1)

Publication Number Publication Date
US6438958B1 true US6438958B1 (en) 2002-08-27

Family

ID=24045194

Family Applications (2)

Application Number Title Priority Date Filing Date
US09/513,943 Expired - Lifetime US6438958B1 (en) 2000-02-28 2000-02-28 Apparatus for reducing heat load in combustor panels
US10/192,576 Expired - Lifetime US6519850B2 (en) 2000-02-28 2002-07-10 Methods for reducing heat load in combustor panels

Family Applications After (1)

Application Number Title Priority Date Filing Date
US10/192,576 Expired - Lifetime US6519850B2 (en) 2000-02-28 2002-07-10 Methods for reducing heat load in combustor panels

Country Status (6)

Country Link
US (2) US6438958B1 (en)
EP (1) EP1132686B1 (en)
JP (1) JP2001248839A (en)
BR (1) BR0100759B1 (en)
CA (1) CA2337311C (en)
DE (1) DE60122954T2 (en)

Cited By (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6557350B2 (en) * 2001-05-17 2003-05-06 General Electric Company Method and apparatus for cooling gas turbine engine igniter tubes
US6655146B2 (en) * 2001-07-31 2003-12-02 General Electric Company Hybrid film cooled combustor liner
US20040035115A1 (en) * 2002-08-22 2004-02-26 Gilbert Farmer Combustor dome for gas turbine engine
US20040250549A1 (en) * 2001-11-15 2004-12-16 Roland Liebe Annular combustion chamber for a gas turbine
US20050050896A1 (en) * 2003-09-10 2005-03-10 Mcmasters Marie Ann Thick coated combustor liner
US20060037322A1 (en) * 2003-10-09 2006-02-23 Burd Steven W Gas turbine annular combustor having a first converging volume and a second converging volume, converging less gradually than the first converging volume
US20080016874A1 (en) * 2004-08-24 2008-01-24 Lorin Markarian Gas turbine floating collar arrangement
US20080134661A1 (en) * 2006-12-07 2008-06-12 Snecma Chamber endwall, method of producing it, combustion chamber comprising it, and turbine engine equipped therewith
US20080145211A1 (en) * 2006-12-19 2008-06-19 Rolls-Royce Plc Wall elements for gas turbine engine components
US20100162716A1 (en) * 2008-12-29 2010-07-01 Bastnagel Philip M Paneled combustion liner
US20130014510A1 (en) * 2011-07-15 2013-01-17 United Technologies Corporation Coated gas turbine components
US10386067B2 (en) * 2016-09-15 2019-08-20 United Technologies Corporation Wall panel assembly for a gas turbine engine
US10807163B2 (en) 2014-07-14 2020-10-20 Raytheon Technologies Corporation Additive manufactured surface finish
US20220299206A1 (en) * 2021-03-19 2022-09-22 Raytheon Technologies Corporation Cmc stepped combustor liner

Families Citing this family (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6672067B2 (en) * 2002-02-27 2004-01-06 General Electric Company Corrugated cowl for combustor of a gas turbine engine and method for configuring same
US6986201B2 (en) * 2002-12-04 2006-01-17 General Electric Company Methods for replacing combustor liners
US6904676B2 (en) * 2002-12-04 2005-06-14 General Electric Company Methods for replacing a portion of a combustor liner
US7802431B2 (en) * 2006-07-27 2010-09-28 Siemens Energy, Inc. Combustor liner with reverse flow for gas turbine engine
FR2914986B1 (en) * 2007-04-12 2015-04-10 Saint Gobain Isover INTERNAL COMBUSTION BURNER
US9810081B2 (en) 2010-06-11 2017-11-07 Siemens Energy, Inc. Cooled conduit for conveying combustion gases
WO2015074052A1 (en) 2013-11-18 2015-05-21 United Technologies Corporation Swept combustor liner panels for gas turbine engine combustor

Citations (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4259842A (en) 1978-12-11 1981-04-07 General Electric Company Combustor liner slot with cooled props
JPS594824A (en) 1982-06-29 1984-01-11 Toshiba Corp Structure of hot gas turbine combustor unit
EP0136071A1 (en) 1983-08-26 1985-04-03 Westinghouse Electric Corporation Varying thickness thermal barrier for combustion turbine baskets
US4628694A (en) * 1983-12-19 1986-12-16 General Electric Company Fabricated liner article and method
US4655044A (en) 1983-12-21 1987-04-07 United Technologies Corporation Coated high temperature combustor liner
US5113660A (en) * 1990-06-27 1992-05-19 The United States Of America As Represented By The Secretary Of The Air Force High temperature combustor liner
EP0493304A1 (en) 1990-12-24 1992-07-01 United Technologies Corporation Integrated connector/airtube for a turbomachine's combustion chamber walls
US5331816A (en) * 1992-10-13 1994-07-26 United Technologies Corporation Gas turbine engine combustor fiber reinforced glass ceramic matrix liner with embedded refractory ceramic tiles
US5553455A (en) * 1987-12-21 1996-09-10 United Technologies Corporation Hybrid ceramic article
US5851679A (en) * 1996-12-17 1998-12-22 General Electric Company Multilayer dielectric stack coated part for contact with combustion gases
US6047539A (en) * 1998-04-30 2000-04-11 General Electric Company Method of protecting gas turbine combustor components against water erosion and hot corrosion

Family Cites Families (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS60238619A (en) * 1984-05-14 1985-11-27 Hitachi Ltd Heat shielding coating structure for burner
JPS62210329A (en) * 1986-03-12 1987-09-16 Hitachi Ltd Ceramic coated heat-resistant material and manufacture thereof
JPH0339821A (en) * 1989-07-04 1991-02-20 Hitachi Ltd Burner
JP2779260B2 (en) * 1990-09-05 1998-07-23 株式会社次世代航空機基盤技術研究所 Gas turbine combustor
US5749229A (en) 1995-10-13 1998-05-12 General Electric Company Thermal spreading combustor liner
JPH11246275A (en) * 1998-03-02 1999-09-14 Senshin Zairyo Riyo Gas Generator Kenkyusho:Kk Heat resistant fiber reinforced composite material and production thereof

Patent Citations (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4259842A (en) 1978-12-11 1981-04-07 General Electric Company Combustor liner slot with cooled props
JPS594824A (en) 1982-06-29 1984-01-11 Toshiba Corp Structure of hot gas turbine combustor unit
EP0136071A1 (en) 1983-08-26 1985-04-03 Westinghouse Electric Corporation Varying thickness thermal barrier for combustion turbine baskets
US4628694A (en) * 1983-12-19 1986-12-16 General Electric Company Fabricated liner article and method
US4655044A (en) 1983-12-21 1987-04-07 United Technologies Corporation Coated high temperature combustor liner
US5553455A (en) * 1987-12-21 1996-09-10 United Technologies Corporation Hybrid ceramic article
US5113660A (en) * 1990-06-27 1992-05-19 The United States Of America As Represented By The Secretary Of The Air Force High temperature combustor liner
EP0493304A1 (en) 1990-12-24 1992-07-01 United Technologies Corporation Integrated connector/airtube for a turbomachine's combustion chamber walls
US5331816A (en) * 1992-10-13 1994-07-26 United Technologies Corporation Gas turbine engine combustor fiber reinforced glass ceramic matrix liner with embedded refractory ceramic tiles
US5851679A (en) * 1996-12-17 1998-12-22 General Electric Company Multilayer dielectric stack coated part for contact with combustion gases
US6047539A (en) * 1998-04-30 2000-04-11 General Electric Company Method of protecting gas turbine combustor components against water erosion and hot corrosion

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
"Physics Part I", by Halliday and Resnick, John Wiley & Sons, Inc., 1966, p. 552. *

Cited By (22)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6557350B2 (en) * 2001-05-17 2003-05-06 General Electric Company Method and apparatus for cooling gas turbine engine igniter tubes
US6655146B2 (en) * 2001-07-31 2003-12-02 General Electric Company Hybrid film cooled combustor liner
US20040250549A1 (en) * 2001-11-15 2004-12-16 Roland Liebe Annular combustion chamber for a gas turbine
US20040035115A1 (en) * 2002-08-22 2004-02-26 Gilbert Farmer Combustor dome for gas turbine engine
US6725667B2 (en) * 2002-08-22 2004-04-27 General Electric Company Combustor dome for gas turbine engine
US20050050896A1 (en) * 2003-09-10 2005-03-10 Mcmasters Marie Ann Thick coated combustor liner
US7007481B2 (en) * 2003-09-10 2006-03-07 General Electric Company Thick coated combustor liner
US20060037322A1 (en) * 2003-10-09 2006-02-23 Burd Steven W Gas turbine annular combustor having a first converging volume and a second converging volume, converging less gradually than the first converging volume
US7093441B2 (en) * 2003-10-09 2006-08-22 United Technologies Corporation Gas turbine annular combustor having a first converging volume and a second converging volume, converging less gradually than the first converging volume
US7690207B2 (en) * 2004-08-24 2010-04-06 Pratt & Whitney Canada Corp. Gas turbine floating collar arrangement
US20080016874A1 (en) * 2004-08-24 2008-01-24 Lorin Markarian Gas turbine floating collar arrangement
US20080134661A1 (en) * 2006-12-07 2008-06-12 Snecma Chamber endwall, method of producing it, combustion chamber comprising it, and turbine engine equipped therewith
US7954327B2 (en) * 2006-12-07 2011-06-07 Snecma Chamber endwall, method of producing it, combustion chamber comprising it, and turbine engine equipped therewith
US20080145211A1 (en) * 2006-12-19 2008-06-19 Rolls-Royce Plc Wall elements for gas turbine engine components
US20100162716A1 (en) * 2008-12-29 2010-07-01 Bastnagel Philip M Paneled combustion liner
US8453455B2 (en) 2008-12-29 2013-06-04 Rolls-Royce Corporation Paneled combustion liner having nodes
US20130014510A1 (en) * 2011-07-15 2013-01-17 United Technologies Corporation Coated gas turbine components
US10113435B2 (en) * 2011-07-15 2018-10-30 United Technologies Corporation Coated gas turbine components
US10807163B2 (en) 2014-07-14 2020-10-20 Raytheon Technologies Corporation Additive manufactured surface finish
US10386067B2 (en) * 2016-09-15 2019-08-20 United Technologies Corporation Wall panel assembly for a gas turbine engine
US20220299206A1 (en) * 2021-03-19 2022-09-22 Raytheon Technologies Corporation Cmc stepped combustor liner
US11867402B2 (en) * 2021-03-19 2024-01-09 Rtx Corporation CMC stepped combustor liner

Also Published As

Publication number Publication date
CA2337311C (en) 2008-08-12
DE60122954D1 (en) 2006-10-26
BR0100759B1 (en) 2009-01-13
BR0100759A (en) 2001-11-06
JP2001248839A (en) 2001-09-14
EP1132686A1 (en) 2001-09-12
US6519850B2 (en) 2003-02-18
US20020178728A1 (en) 2002-12-05
EP1132686B1 (en) 2006-09-13
CA2337311A1 (en) 2001-08-28
DE60122954T2 (en) 2007-05-03

Similar Documents

Publication Publication Date Title
US6438958B1 (en) Apparatus for reducing heat load in combustor panels
US6250082B1 (en) Combustor rear facing step hot side contour method and apparatus
EP2549188B1 (en) Insert for a gas turbine engine combustor
US10041676B2 (en) Sealed conical-flat dome for flight engine combustors
EP1010944B1 (en) Cooling and connecting device for a liner of a gas turbine engine combustor
EP3008387B1 (en) Conductive panel surface cooling augmentation for gas turbine engine combustor
EP3047128B1 (en) Controlled variation of pressure drop through effusion cooling in a double walled combustor of a gas turbine engine
US20200025376A1 (en) Gas turbine engine combustor with tailored temperature profile
US20180230602A1 (en) Coated combustor panel shell for a gas turbine engine combustor
US20160320060A1 (en) Thermal mechanical dimple array for a combustor wall assembly
US6986201B2 (en) Methods for replacing combustor liners
EP2573464B1 (en) Combustion sections of gas turbine engines with convection shield assemblies
US10935243B2 (en) Regulated combustor liner panel for a gas turbine engine combustor
US7360364B2 (en) Method and apparatus for assembling gas turbine engine combustors
EP3084307B1 (en) Dilution passage arrangement for gas turbine engine combustor
EP1132687A1 (en) Methods and apparatus for reducing thermal stresses in an augmentor

Legal Events

Date Code Title Description
AS Assignment

Owner name: GENERAL ELECTRIC COMPANY, NEW YORK

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:MC CAFFREY, TIMOTHY P.;LASTRINA, FRANK A.;MONTY, JOSEPH D.;AND OTHERS;REEL/FRAME:010874/0132

Effective date: 20000406

STCF Information on status: patent grant

Free format text: PATENTED CASE

CC Certificate of correction
FPAY Fee payment

Year of fee payment: 4

FPAY Fee payment

Year of fee payment: 8

FPAY Fee payment

Year of fee payment: 12